Aircraft



Jan. 22, 1935. -r Re. 19,438

AIRCRAFT Original Filed Dec. 30, 1927 6 Sheets-Sheet l INVENTOR:

Jan. 22, 1935. H. H. PLATT 7 Re. 19,438

' AIRCRAFT L I J H J 57 V SECTION 01v LIA/5.315.

INVENTOR:

HAW/Z mm H. Pm rr, I

H. H. PLATT Jan. 22, 1935.

AIRCRAFT Original Filed Dec. 50, 1927 6 Sheets-Sheet 4 t 4 INVENTOR:H/YV/L/YND li P0777;

Jan. 22, 1935.

H. H. PLATT AIRCRAFT Original Filed Dec. 30, 1927 6 Sheets-Sheet 5INVENTOR: HPL/Yrr,

H. H. PLATT Jan; 22, 1935.

AIRCRAFT Original Filed Dec. 30, 1927 6 Sheets-Sheet 6 R O T N E V NReissued Jan. 22, 1935 UNITED STATES AIRCRAFT Haviland H. Platt, NewYork, N. Y., assignor to Frederick W. Wilkening, Philadelphia, Pa.

Original No. 1,795,501, dated March 10, 1931,

Serial No. 243,522, December 30, 1927. Application for reissue November18, 1932, Serial 22 Claims. (Cl. 244-41) My invention relates to aheavier-than-air type of aircraft and it relates more particularly to anaircraft of this type, which is capable of forms and degrees ofnavigability unobtainable 5 with the standard types of aircraftheretofore devised.

My invention further relates to a heavier-thanair type of aircraftwhich, by virtue of universal navigability inherent in the structurethereof, is of greatly increased utility over the present standard formsof aircraft, because it is capable of taking to flight and landing underad- I verse conditions, of which present standard forms of aircraft areincapable, and because it is also capable of a universal control ornavigability in mid-air, which aircraft of the present standard formsare likewise incapable of.

It is well-known that with standard aircraft structures, particularly ofthe heavier-than-air type, certain limitations are imposed upon the useof the aircraft by the size and character of thelanding field with itssurrounding obstructions, as well as by the degree of visibility when inflight. Thus, with any of the many present forms of standard aircraftand particularly so with the larger sizes of aircraft when heavily orfully loaded, a relatively large field is required with a uniformly evenand firm surface, in order for the aircraft either to take off withsafety, or to land with safety.

-This requirement as to the size and quality of the landing field isaccentuated by the increased size or increased loading of the aircraft,and at best, a field of considerable magnitude is required at presentfor heavier-than-air aircraft in order for the same to land or take off,which requirement greatly reduces the ultimate utility or practicabilityof this type of aircraft, aside from the natural hazards attendant upontaking off or landing machines of this type at high speed. Aside fromthe limitations of the presentstandard aircraft in taking off andlanding, they are further limited in their ultimate utility by the factthat they require a relatively high air speed in order to maintain themaloft, which inherently increases the hazards of collision with otheraircraft or with ground objects or obstructions, particularly so withpoor visibility.

By my novel aircraft construction, the above limitations of theheavier-than-air type of machine are eliminated by the followingfeatures of navigability inherent in this new construction:

(1) Capacity for vertical as well as horizontal propulsion or propulsionunder power, in any direction intermediate the vertical or horizontal.

Due to this, the aircraft can take off without afield of anyconsiderable size, and can be maintained aloft without any lateral orhorizontal velocity or with any small horizontal velocity.

(2) Ability to change the direction of propulsion or the angle of thrustin any direction while in flight, in a vertical ,medium plane of thecraft.

Hence the craft is not only capable of vertical as well as. horizontalpropulsion and capable of being maintained stationary in air, but byproperly changing and directing the angle of thrust or direction ofpropulsion the speed of the craft in any direction, particularly in theforward direce tion, can be decelerated and in fact, the velocity of themachine brought to zero in that direction, in a manner very similar tothe braking or stopping of road vehicles.

This latter feature of navigability which can be resorted to in mid-air,automatically decreases the hazard of collision at high speed tosubstantially the same degree as is now inherent in road vehicles.

In addition to the above features of navigability under power, the craftof this novel construction is also capable of certain other features of'navigability, particularly in landing without power. Thus, for instance,it is capable of relatively low vertical speeds in either inclined orvertical descent without power, thus increasing the safe navigability ofthe craft under emergency conditions. I

With the above ends in view my novel invention consists of z (1) Theprovision of motor propelled rotatable aerofoils upon the body of thecraft which perform the dual primary functions of propulsion and lift,performed in the present standard form of aerofoils by the so-calledpropeller and wings, respectively.

(2) The provision of a number of such aerofoils disposed horizontallyand transversely of for varying and controlling, at the will of theoperator, the direction of the thrust thus produce'd by" the rotor.

(4) Means for varying and controlling, at the -will of the operator, themagnitude of the thrust thus produced. r v The provision of a pair ofsuch rotors, one on each side of the machine, and means for eitherjointly or independently varying and controlling both the direction ofpropulson or angle of thrust, as well as the magnitude of suchpropulsion or thrust in each of the two rotors, thereby to attainlateral navigability such as turning, banking and the like. I

(6) The invention also consists in so relating the propelling forces andthe center of gravity of the fuselage of the aircraft, with respect tothe axis of the rotors, that the reactionupon the fuselage, of thedriving power transmitted to the rotors, will tend to raise or elevatethe rear or tail end of the fuselage to a suitable extent, and so thatupon the diminution or cessation of the power delivered to the rotors,the fuselage will tend to nose-up" to a suitable extent, therebyfacilitating the taking-off, as well as the landing of the aircraft.

In the accompanying drawings, I have illustrated in a general way, aform of construction embodying the above structural and functionalcharacteristics, although without any attempt to show any detailedconstruction.

Thus, Figure 1 represents a top plan view of an aircraft of my novelconstruction showing the fuselage, partly in section and showing thepair of transversely disposed rotors (one partly broken away).

Figure 2 represents a side elevation of an aircraft of thisconstruction, illustrating further the general relation of the rotor tothe fuselage.

Figure 3 represents a section on line 3--3 of Figure 2 (with one of therotors not shown) illustrating the method of drivingthe propeller orrotor and the method of controlling the angle and magnitude of thethrust of the rotor, from within the fuselage.

Figure 4 represents a partial side elevation, on

an enlarged scale, of 'the ae'rofoil or blade deflecting arms of therotor and the driving and controlling means within the rotor.

Figure 5 represents a section on line 5--5 of Figure 4, illustratinggenerally the method of driving or propelling, and the method ofactuating the blade deflecting arms and also the manner of varying orcontrolling the movements of the deflecting arms, thereby to control theangle and magnitude of the resultant thrust of the rotor as a whole. I

Figure 6 represents a section on line 66 of Figure 3, on an enlargedscale, illustrating the means within the fuselage for controlling theangle as well as the magnitude of the thrust of each of the two rotorseither jointly or independently of each other.

. blade angles.

Accordingly, Figure 8 is a diagram of rotor and controls in the straightor full-speed-ahead position, under power.

Figure 9 represents a similar diagram of rotor and controls for verticalascent or descent or hovering, under power.

Figure 10 represents a similar diagram of rotor and controls in aposition intermediate to that shown in Figures 8 and 9, for an inclinedascent under power.

Figure 11 represents a similar diagram of rotor and controls with thedirection of propulsion or angle of thrust shifted into a rearwardreverse direction or opposed to the direction of travel, fordecelerating the machine.

Figure 12 represents a similar diagram of rotor and controls showing thesame in a position for vertical descent, without power, and with therotors free to revolve.

Figure 13 represents a similar diagram of rotor and controls with theangle of thrust directed downwardly for running the engine under load onthe ground, for the purpose of warming the engine.

In Figures 1, 2 and 3 of the accompanying drawings, the fuselage of theair craft of my novel construction is designated generally by thenumeral 15. The fuselage may be of any suitable shape and construction,as may be required by the necessary strength, the .wind resistance, andgravity forces.

It will be noted that the aircraft of my novel construction lacks thefixed lifting surfaces or what are commonly termed wings, whichinvariably form an indispensable part of the conventional aircraft nowin use, and it likewise lacks the conventional form of propeller.

Instead, the aircraft of my novel construction is provided with, whatmay arbitrarily be termed, rotors. The rotors, designated generally bynumerals 16 and 17, are disposed symmetrically oneach side of thefuselage, upon a common horizontal axis extending transversely of thefuselage.

Each rotor comprises a plurality of aerofoils or what may be termedblades 18. vThe number of blades in a rotor may be varied between somepractical limits, depending upon the size of the rotors as well as uponcertain other factors. In the illustration of my novel aircraft shown inthe accompanying drawings each rotor containsfour blades or aerofoils;that number having been found to give satisfactory results. For somepurposes, rotors of three blades as well as possibly five and sixblades, may also be desirable.

The two rotors 16 and 17 are supported and driven by a common axialhorizontal shaft 19 extending transversely through the fuselage and maybe suitably journalled in a central bearing 20 as well as in a pair ofopposed terminal bearings 21, carried by and preferably forming part ofa driving shaft and gear housing 22, near the top of the fuselage.

The shaft and gear housing 22, being rigid with the fuselage 15,projects outwardly from each side of the fuselage, as seen particularlyin Figures 1 and 3, in gradually tapering form, so as to give thedriving shaft 19 the necessary rigid support near the point of the load,to wit,- -near the outer extremities thereof.

The opposed ends of the main driving shaft 19 extend through thebearings 21, in each end of the cent to the bearing sleeves 23, suitableflanges 24 are fixedly secured. A'driving arm spide1125 is provided atthe end of the main driving shaft 19, having a hub 26 and a number ofsimilar radial driving arms 27, extending tothe number of blades"comprising the rotor. A shaft exten sion 28, having a correspondingflange 29 is secured to each of the opposed ends of the main shaft 19 aswell as to the corresponding driving arm spider 25, by means of a seriesof bolts 30 extending through the two flanges 24 and 29, on

the main shaft 19 and shaft extensions 28, respectively, and extendingalso through the hub 26 of the arm spider 25.

Each of the blades 18 is pivotally secured to the corresponding drivingarm 27, at a point substantially midway of its length, and with the axisof the pivot being substantially at the center of gravity of thecross-section of the blade, which latter depends upon the particularcross-section of the blade.

The blades 18 may be of any suitable internal construction, althoughpreferably formed of metal. Thus, the blades 18 are preferablythinwalled, metallic shells, hollow within, except for bracing webs ofmetal to give it the necessary rigidity and strength, and are preferablyformed of some of the well-known aluminum alloys. The method of pivotingthe blade 18 to the driving arm 27 (as well as other supporting anddeflecting arms to be described hereinafter) is not shown in detail inthe accompanying drawings, and any suitable pivot construction may beemployed for the purpose.

ill

In order to give each blade 18 the necessary resistance to centrifugalforces, while the rotor is in motion, and in order to give the rotors l6and 17 as a whole, the necessary rigidity, the blades 18 are furthersupported at or near their opposed ends, by corresponding arms 31 and32. respectively, to which the blades are similarly pivotally secured atthe center of gravity of their cross-sections and in axial alignmentwith the pivotal support of the main driving arms 2..

Each set of supporting arms 31 is formed integral with, and is carriedby a. common annular hub or ring 33 which loosely surrounds the corresponding tapering shaft and gear housing 22.

The outer supporting arms 32, on the other hand, are also preferablyformed integral with a central hub 34, which is secured to a terminalflange 35 on the outer ends of the shaft extensions 28, by means ofbolts 36.

By this means, each blade is supported at its center of gravity againstthe centrifugal force, as

well as against air pressure, at three points along its entire length,while the driving support, through the arms 2'7, is positionedsubstantially midway of its length.

As pointed out hereinabove, the rotors, consisting of a, plurality ofpivotally mounted aerofoils or blades perform two separate and distinctprimary functions; one is the propeller function and the other is thewing or lift function. To perform these two functions the rotor or moreparticularly each blade of the rotor, must be presented to the airstream at various angles as it travels around in a substantiallycircular path about the axis of the shaft 19. f

The arrow 37 in FiguresZ and 8 to 13 inclusive indicate the direction ofrotation of the rotor, as well as the path of the blades comprising thesame.

Inorder to vary the angle of .each blade at different parts of itstravel in the circular path,

each blade is connected to a common, normally stationary eccentric, bymeans of an eccentric strap 38 and a deflecting rod 39, which is rigidand preferably integral with the eccentric strap and the free end ofwhich is pivotally secured to the blade 18 in advance of the supportingpivot, pref erably near the leading edge thereof.

By this means, each blade is deflected in succession, between twoopposed extreme limiting angular positions, as it passes through eachrevo lution or as it travels through the circular orbit of the blades.The positions in the circle, of the two limiting angles of the blades intheir circular orbit, is determined by the direction of the eccentricityof the eccentric, while the magnitude of deflect. on of the blades,either from their median position or between their two extremepositions, is determined by the amount of the eccentricity.

The direction of the eccentricity therefore determines generally thedirection of the forces produced by the action of the blades upon theair stream, while the amount of the eccentricity determines generallythe magnitude of the forces so produced.

In order to vary the direction and the amount 0r magnitude of theeccentricity, and thus to obtain the necessary navigability of thecraft,the eccentric straps 33, instead of being merely supported upon anormally stationary eccentric, are

instead supported or carried by a dual eccentric, 11

both constituent eccentrics of which are normal- 1y stationary butmovable with respect to the fuselage as well as with respect to eachother, for adjustment or control purposes.

Thus, upon referring particularly to Figures 3, 5 7

4 and 5, it will be-noted that the eccentric straps 33 are rotatablymounted upon an outer eccentric 40, which in turn is rotatably mountedupon an inner eccentric 41. The inner eccentric 41 in turn is rotatablymounted upon the concentric bearing sleeve 23 which is rigidly aflixedto, or is integral with the shaft and gear housing 22 and terminal mainbearing 21.

The inner eccentric 41 is integral with, or is rigidly and fixedlysecured to a sleeve 42, to the opposed end of which a pinion 43 isfixedly secured. A pinion 44 is rotatably mounted upon the sleeve 42,and carries a flange 45. The flange 45 is provided with a generallyradial slot 46 which engages a laterally extending pin or projection 47flxed in the outer eccentric 40. By this means, the outer eccentric maybe revolved in either direction for adjustment or control purposes, bymerely revolving the concentric pinion 44; the motion between theconcentric pinion 44 and flange or disc and the eccentric 40 andeccentric pin 47, being effected by the engagement between the slot 46and the pin 47.

The inner eccentric 41, on the other hand, may be revolved in eitherdirection for adjustment or control purposes, by merely revolving theconcentric pinion 43.

In order to maintain the two eccentrics 40 and 41 in any desiredrelative position with respect to the fuselage and with respect to eachother, a pair of concentric control shafts 48 and 49 are provided oneach side of the fuselage, corresponding to each of the rotors 16 and17. The inner control shafts 49 are journalled within the outer tubularcontrol shafts 48, while the latter are journalled in suitable bearings50 in the fuselage,

inner control shafts 49, pinions 51 are fixedly secured, which areconstantly inmesh with the pinions 44 which actuate the outer eccentrics40 To the inner ends of each of the inner control shafts 49, sprocketwheels 52 are secured, over which the sprocket chains 53 pass.

To the outer ends of each of the tubular control shafts 48, similarpinions 54 are fixedly secured,

' compartment where they ,pass around the corresponding sprocket wheels'7 and 58. respectively,

forming part of the control mechanism within the pilot's compartment ofthe fuselage.

The pilot's control mechanism is shown in detail particularly in Figures3, 6 and '7, while the corresponding control mechanisms in the rotorsare illustrated in detail particularly in Figures 3, 4

and 5. Each of the pair of opposed, sprocket wheels 5'7 is carried by atubular shaft 59 journalled in corresponding and similar bearings 60which are rigidly carried upon brackets 61 fixed to the fuselage. To theinner and juxtaposed ends of the pair of tubular shafts 59, acorresponding pair of bevelled gears or bevelled gear sectors 62 arefixedly secured in a manner shown particularly in Figure 7. Adifferential bracket 63 is journalled or rotatably mounted upon one ofthe tubular shafts 59 by means of the bearing 64 thereof. A lower andright angular bearing 65 of the differential bracket 63 in turn carriesthe differential or steering shaft 66, to the inner end of which abevelled pinion 6'7 is fixedly secured, which in turn is in meshwitheach of the two juxtaposed bevelled gear sectors 62, therebydifferentially interconnecting the same at all times. The outer end ofthe shaft 66 carries a steering wheel or other handle member 68.

By deflecting the differential or steering shaft 66, as well as thebracket 63, in either direction 69 or '70, indicated in Figure 6, theouter eccentric 40, of each of the two rotors on either side of thefuselage, will be revolved in either one of two opposed directions, tothe same degree or extent. On the other hand, by revolving the wheel orhandle member 68 and the shaft 66, in either direction '71 or '72(indicated in Figure 3) the outer eccentric 40 of each of the tworotors, will be defiected or displaced in opposed directions. Thus,

by the deflection of the shaft 66 in either directions 69 and '70 or therotation of said shaft 66 in either directions '71 or '72, a uniform ora differential adjustment or control of the outer eccentrics 40 in therotors l6 and 1'7 is obtainable.

Each of the sprocket wheels 58, on the other hand, is carried bycorresponding shafts 73 which extend through the hollow tubular shafts59 and project beyond the same at each end. To the inner end of the twoshafts 73, lever arms '74 and '75 respectively, are secured, which carryarcuate extensions '76 and '77, to which corresponding handles '78 and'79 are secured. The arcuate extensions '76 and'7'7 areparallel to eachother and are arranged frictionally or positively to interlock with eachother at will. This friction or positive interlocking means may be anysuitable detent; catch or the like, which is not shown in the drawings.An arcuate sector 80 is rigidly carried by the bearing65 of the bracket63, and parallel to the arcuate lever extensions '76 and '77. Suitablefrictional or positive inter ldcking means are also provided betweeneitherv one of the arcuate lever extensions '76 and '77 (or both) andthe sector 80, which is stationary with respect to the differentialbracket 63, whereby either one or both, of the levers '74 and 75 andhence the corresponding pair of shafts '73 may be interlocked with thedifferential bracket 63 and hence interlocked with the pair of tubularshafts 59, in any desired position intermediate of suitable limits.

By such an interlocking of either one or both of the lever handles '78and '79 and the levers '74 and '75 with the differential bracket 63,either one or both of the corresponding inner eccentrics 41 in the tworotors respectively, may be revolved or controlled in unison with theouter eccentrics 40. I

The operation The operation of the aircraft, or more particularly, thecontrol and navigation thereof can best be understood upon a furtherreference to Figures 8 to 13 inclusive. In these figures the rotor andcontrols within the fuselage are rep resented diagrammatically. InFigure 8, one of the eccentric straps 38, and the deflecting arm 39 areshown, while in Figures 9 to 13 inclusive the deflecting arms 39 aremerely represented by dotted lines. In these figures the inner and outereccentrics are represented by corresponding circles designated as 41 and40 respectively. The fixed or driving center of the rotor, that is, theaxial center of the driving shaft 19 and bearing sleeve 23, isrepresented by the intersection of vertical and horizontal referencelines, represented as dot and dash lines. The dotted circle 44 in eachof these diagrammatic figures represents the pinions 43 and'44,concentric with the main driving shaft 19 and bearing sleeve 23, whilethe dotted circle 51 in these figures represents.the pinions 51 and 54,by means of which the former pinions are driven for purposes ofeccentric control. The dotted line 56 represents the sprocket chains 53and 56, while the solid circle 58 represents the two sprocket wheels 58and 5'7 within the fuselage. The manual control mechanism is also markedby appropriate reference characters, correspond ing generally to thedetailed view of this mechanism, as shown in Figures 3, 6 and '7.

In each of these diagrammatic figures, moreover, the eccentricity ofeach of the two eccentrics is represented by corresponding heavy andlight lines. Thus the heavy line extending from the fixed center to thecenter of the eccentric circle 41, represents the eccentricity of theinner eccentric, while the light line extending from the center of theinner eccentric circle 41. to the center of the outer eccentric, 40represents the eccentricity of the outer eccentric.

It will be observed here, that-while thesupporting centers 81 of theblade 18, which are pivotally secured to the driving arms 2'7 andbracing or supporting arms 31 and 32, revolve in a true and'flxed circleabout the fixed center (at the intersection of the twodot and dashreference lines), so too, every other part of the blade 18 likewiserevolves about a true circle, though eccentric with the first circle.The center of rotation of each of the various parts of the blade 18 isdetermined by the resultant or effective or total eccentricity of theinner and outer eccentric. The circle 82 in each of these diagrammaticproximately the center of gravity of the crosssection of the blade),while the circles 83 represent the path of travel of the deflectingpivot 84 of the blade 18, and hence represent generally the path oftravel of the leading edge of said blade.

While the circle 82 is fixed, and has its center at the center of thedriving shaft 19, the circle 83 may be varied at will, by means of theeccentric control.

The arrow 85 represents the direction of rotation of the rotor, whilethe arrow 86 represents the forward direction of the aircraft.

As pointed out heretofore, the magnitude of the force produced by therotor is determined by the maximum deflection or angularity of the bladeangle during each revolution, which in turn is determined by the amountor magnitude of the eccentricity. This is indicated in the diagrammaticviews, by the distance between the fixed center (intersection of the twodot and dash reference lines) and the center of the outer ec centriccircle 40 (the free. end of the light eccentric line).

As also brought out hereinabove, the direction of the force produced bythe rotor is determined, on the other hand, by the relative point in thecircular path of the blades at which they are deflected to their maximumdeflection as determined by the direction or angle of the resultant ofeffective eccentricity. This latter is indicated in the diagrammaticviews by the imaginary line which passes through the fixed center andthe center of the outer eccentric circle 40.

Under normal condition, straight flight in any direction, eitherhorizontally, straight away or on an upward ascending incline, or in avertical rise, is effected by uniform setting of both rotors. Theunequal setting of the two rotors is required only under one of threeconditions; one is to balance the aircraft when it becomes necessary todo so for any reason whatsoever, that is, when it is necessary to exertunequal forces on the two sides of the aircraft in order to maintain iton an even keel; second, to change the direction of flight or to turn,and to bank the aircraft simul taneously; and third, to effect adisplacement of the aircraft sideways.

The navigation of the aircraft under these last three conditions, by anunequal setting of the two rotors, will be discussed hereinafter. Thusthe reference to the diagrammatic views of Figures 8 to 13 inclusive, ismade primarily with respect to an equal or uniform setting of the tworotors for flight solely in straight lines.

' In Figure 8 the rotor and controls are shown in a position for fullspeed ahead. For full speed ahead, at a uniform altitude, the angle ofthe effective eccentricity or the direction of thrust of the rotor isinclined upwardly above the horizontal, just sufiiciently to maintainthe craft at uniform altitude. This depends upon the loading of thecraft and upon the forward speed. The angle or maximum deflection of theblade 18 in turn is so adjusted as to present the blade at the mostefficient angle of attack with respect to the relative air stream. Thismagnitude of the eccentricity or of the rotor thrust, depends amongothers, also upon the loading and upon engine power, as well as upon thespeed of the aircraft.

It should be observed here that each blade'performs the two distinctfunctions of propulsion" and lift, generally at two different parts of ipath of travel in a circle.

Thus, when the blade is at the upper part of *the circular orbit itexerts a propelling force,

while when the blade is at the lower part of its circular orbit andtravels in the same direction as the fuselage, it. exerts a liftingforce. The lifting force thus produced by the blade is greatly in excessof the lifting force that would ordinarily be produced by a fixed wingof the same projected area because of the fact that the blade whenacting as a lifting wing" in the lower part of its circular orbit, istraveling in the same direction as the fuselage and the speed relativeto the fuselage is as great or greater than the air speed of thefuselage, and hence the resultant airspeed of the blade in the lowerpart of its circular orbit is double the speed, or more, of the airspeed of the fuselage. v

For this reason, the aircraft of this construction can be maintainedaloft with aerofoil sur-" faces of an aggregate projected area less thanthe fixed area required to lift the same weight at the same speed of\theaircraft.

After the inner and outer eccentrics have been set with respect to eachother by locking the two levers 74 and '75 to the differentialbracket'63, in the desired position, and'thereby setting the magnitudeof the eccentricity, with the condition of flight as above outlined andindicated in Figure 8, the altitude of the aircraft may be raised orlowered at will, by merely changing the angle of thrust to the desireddegree in the desired direction, without however, necessarily changingthe magnitude of the thrust or the magnitude of eccentricity.

Thus, while the craft is in straight-away flight full-speed-ahead, thecraft can be put into an inclined ascent or descent, primarily byraising or lowering the angle of the resultant or effectiveeccentricity. This is accomplished merely by deflecting upwardly ordownwardly, in either of the directions indicated by the arrows 70 and69, respectively, (Figure 8 and Figure 6) the differential shaft 66, aswell as the differential bracket 63, without, however, turning the wheel68.

Thus it will be observed that the raising of the shaft 66 in thedirection of the arrow 70, will raise the angle of the total oreffective eccentricity by revolving the two eccentrics 40 and 41 in thesame direction and to the same extent, thereby retaining the magnitudeof the resultant eccentricity unchanged, By deflecting the shaft 66downwardly in the direction of the arrow 69, the reverse takes place.Thus the inner and outer eccentrics 41 and 40 respectively, are bothrevolved in a counter-direction to the same degree, thereby lowering theangle of eccentricity, without, however, changing the magnitude thereof.

The diagram of Figure 9 represents generally the condition of flight ofthe craft in a straight vertical ascent. It will be observed that toattain this condition of flight, the direction of the resultanteccentricity or the direction of thrust is made substantially verticalin an upward direction by deflecting the operating shaft 66 in1thetdirection of the arrow 70 to the desired ex- Since the speed ofvertical ascent is necessarily considerably less than the forward speedof the craft, it is necessary, for the best efliciency, to

' reduce the blade angle or the magnitude of the a direct retardingforce upon the aircraft, thereresultant or effective eccentricity. Thusforwar- -tical ascent, in addition to deflecting the operating orcontrol shaft 66 in an upward direction, as shown in Figure 9, the twooperating handles 78 and 79, which control the-inner eccentrics 41; mustbe brought nearer to the shaft 66, thereby to revolve the innereccentrics to the desired extent with respect to the outer eccentrics.This is also indicated in Figure 9. Y

The diagram of Figure 10 shows the craft in an upward inclined ascentunder power. The rotors and controls for this condition of flight areset in a position intermediate of those shown in Figures 8 and 9,respectively. Thus theldirection of the resultant or efifectiveeccentricity or the direction of thrust of "the rotor is inclinedupwardly at an angle greater than that required to maintain the craft ata uniform altitude. To compensate for the reduction of air speed, due tothe upward inclined ascent, the magnitude of the eccentricity and hencethe angle of the blade is reduced. Accordingly the control shaft 66 isdeflected upwardly to a position intermediate of that shown in Figures 8and 9, while the inner eccentric control handles 78 and 79 are moved toa position withrespect to the control shaft 66 intermediate of theposition thereof, shown in Figures 8 and 9, thereby to set the magnitudeofthe eccentricity to a quantity intermediate of the maximum quantityshown in Figure 8, and the lower quantity shown in Figure 9.

In Figure 11 I have shown the method of control of the aircraftwhen itis desired to decelerate the speed of the craft, let us say, in aforward direction. This feature of the control of the aircraft isanalogous to the braking of road vehicles.

This is accomplished by merely changing the angle of the resultanteccentricity or the direction of thrust of the rotor, from a position,for instance, that shown in Figure 8 (full speed ahead) to a positionshown in Figure 11, that is, in a somewhat rearward direction or back ofvertical. This rearward deflection of the angle of thrust of the rotormust be accompanied, however, by a decrease in the amount ofeccentricity or a decrease in the blade angle, which in turn isaccomplished by bringing the inner eccentric control handles '78 and 79nearer to the control shaft 66. The effect of this control is to exertby checking its speed. As the speed of the craft is decreased by thisoperation, the direction of eccentricity or the direction of thrust isagain brought forward towards the vertical, so as to maintain thedesired altitude of the craft, notwithstanding the reduction in speed,or possibly notwithstanding the bringing of the craft to a standstill.

It will be seen from the foregoing, that the aircraft is navigable underpower either in a straight away flight or an upward or a downwardinclined ascent or descent, as well as a vertical ascent or descent, atpractically any desired horizontal or vertical speed, thereby renderingthe craft universally controllable under power and capable of landingand taking off without any limitations as to size or character oflanding field.

It should be observed, however, that the aircraft of this novelconstruction is not limited in navigation to control under power. It islikewise capable of an inclined or a vertical descent without power, aswould be occasioned by failure of the engine. Under such emergencyconditions the main driving shaft 19, and hence the rotors 16 and 17 aredisconnected from the engine or source of driving power, by means of aclutch 87 (shown only conventionally) which may be operated by anysuitable handle 88 in proximity to the pilot quarters. power from thebevelled gear 89 to the main driving shaft 19 which extends transverselyof the aircraft and carriers at its two opposed ends, the two rotors 16and 17, respectively. The source of power is representeddiagrammatically in Figures 1 and 2, as the internal combustion en-'gine 90, suitably connected with the main driving shaft 19, through thebevelled pinion 91.

In order to navigate the craft without power, that is, under emergencycondition, with the source of power 90 disconnected from the maindriving shaft 19, and with the rotors l6 and 1'7 therefore free torevolve in unison with each other, a third function of the rotors isutilized. Under thiscondition of the flight, the rotors are maintainedin rotation by the windmill action -of the air stream upon the rotors,as the craft passes through the air by the force of gravity.

Thus, for an inclined or gliding descent without power, the blade angleand the direction of thrust or the angle of the resultant eccen--tricity, is set in substantially the position shown in Figure'8 for fullspeed flight ahead, under power. For a vertical descent without power,on the other hand, the eccentricity is set substantially as shown inFigure 12. Under'each of these two conditions of free flight, as well asany intermediate position, the angle of the resultant eccentricity, orwhat may be termed the normal direction of thrust of the rotor (underpower,) is set substantially in line with the direction of travel of theaircraft or substantially in line with the air stream. The effect ofthis is to produce a windmill action upon the rotor, by the "air stream,and thereby to maintain the rotor in rotation.

In a free inclined descent the rotors are maintained in rotary motion bythe action of the air stream upon the blades, when at a certain part ofthe path of their travel, while the same blades when at a different partof their circular path of travel, will be presented to the same airstream at an angle so as to produce a reaction opposed to the force ofgravity, thereby checking the free downward or vertical component oftravel of the aircraft. Here too, the blades in their lower positionsact as gliding surfaces.

In a free vertical descent, the blades act as gliding surfaces both intheir lower and upper positions, that is, when moving transversely ofthe direction of travel of the craft.

It will be observed that for an inclined free gliding travel, with theaircraft set substantially in the condition shown in Figure 8, themagnitude of the eccentricity is considerably greater than the magnitudeof the eccentricity for a free vertical descent, as shown in Figure 12.It is estimated under either one of these two conditions of the free orgliding flight, either in an inclined or a vertical descent orintermediate angles of descent, the vertical speed 'of the craft can bekept down to a minimum which will afford a safe landing, while thehorizontal speed The clutch 87 transmits the of the aircraft is, ineffect, pivotally suspendedfrom the generally horizontal axis of therotors,

and the power delivered from the primary power source, such as theengine or the like, to the rotors, creates a torque reaction upon thefuselage, tending to rotate the fuselage in .a direction opposite to therotation of the rotors, and about the same axis. The center of gravityof the fuselage (the entire structure minus the rotors and rotor shaft)is so located with respect to the axis of the rotors, that if thefuselage were freely suspended from the axis of the rotors, the fuselagewould assume a suitably inclined position with its nose-end up, and withits tail-end down. This is the position which the fuselage thereforetends to assume in all :powerless or free-flight positions, and thelanding wheels 96 and the tailskid 103 are so positioned with respect toeach other and with respect to the fuselage that they will be ingenerally horizontal alignment with each other. when the rotors areeither auto-rotated by the air passing therethrough (as for instance, inan inclined glide or vertical descent without power or with the enginedisconnected) or are merely idling or rotating without substantial powerfrom the engine, that is, when the fuselage is nosed up. As power isapplied to the rotors, the torque reaction on the fuselage tends to liftup the tail-end of the fuselage into the position shown in Figure 2, andthis torque reaction on the fuselage, of the power delivered to therotors, is counterbalanced by the weight of the fuselage below the axisof the rotors, as well as by the vertical control which may be derivedfrom the control surfaces 94 while in forward flight.

Thus, the position assumed by the ship in normal forward flight underpower would be the general position shown in Figure 2, while if,

the power were shut off or diminished to idling while in flight, thenose would tip up and the tail .would be deflected downwardly, becauseof the elimination of the torque reaction upon the fuselage. Thisnosed-up position of the fuselage is suitable for gliding and landing,while the horizontal position shown in Figure 2 is suitable for forwardflight and for taking off the ground, and is attained through theapplication of power to the rotors.

In Figure 13, I have illustrated the setting of the rotors and controlswhen it is desired to warm up the engine in preparation for flight. Thusinstead of anchoring the aircraft as is now necessary while warming upthe engine, it is merely necessary to deflect the angle of theeccentricity or the direction of thrust in a downward verticaldirection, that is, toward the ground. The engine and the rotors maythus be operated indefinitely with the aircraft stationary on theground.

The foregoing is a general outline of the methods of control, forflights in substantially straight direction.

In order to change the direction of a craft while in flight, that is toturn about, a differ ential control of the two rotors 16 and 17 is resorted to.

It will be observed that by deflecting the shaft 66 as well as thehandle wheel 68 in either direction 69 or 70, without, however, turningthe wheel 68, the rotors 16 and 17 are controlled uniformly.

If, however, instead of deflecting the shaft 66 and differential bracket63, the wheel 68 is revolved in either direction 71 or 72 (Figure 3)then the two hollow tubular control shafts 59,

(in the fuselage) and consequently the two control shafts 49 of therotors, and the corresponding two outer eccentrics 4 will be affecteddiffer-entially.

Thus, depending upon the direction of rotation of the wheel 68 the angleof the eccentricity of the eccentric 40 in either one of the two rotors16 and 1'7 will be deflected upwardly with respect to the angle ofeccentricity of the inner eccentric 41, while the same eccentricity inthe other rotor will be deflected downwardly to like extent.

By this operation the resultant eccentricity in the two rotors ischanged both as to direction and as to magnitudeQand hence the thrust ofthe two rotors l6 and 17 is altered differentially both as to directionand magnitude.

It will be observed moreover, that by differentially varying thepositions only of the two outer eccentrics 40, as for instance, in astraight away flight indicated in Figure 8,the resultant eccentricity inone rotor is raised both as to angle and as to magnitude, while in theother rotor the angle of the eccentricity is lowered and the magnitudeis decreased correspondingly. The effect of this is not merely a turningof the craft, due to an unequal thrust exerted by the two rotors, butalso a simultaneous and automatic banking. of the craftdue to a raisingof the angle of the eccentricity or the direction of thrust on theoutside of the turn and a lowering of the same on the inside. c v

The differential control of the inner eccentrics centric handles '78 and79, is utilized primarily for balancing purposes or for stabilizing theaircraft.

In Figures 2, 3 and 8 to 13 inclusive, the pilots seat is designatedgenerally by the numeral 92, while an auxiliary seat shown also indotted lines in Figure 2 is designated by the numeral 93.

1 Since the entire navigation and control of the aircraft is dependentsolely upon the uniform or differential control of the two rotors 16 and1'1, the aircraft of this novel construction does not require anyauxiliary controls, such as ailerons, or vertical or horizontal rudders.

At the tail end of the fuselage 15, a pair of fixed horizontalstabilizers 94 maybe provided of suitable area. Likewise, a stationaryor rigid vertical stabilizer 95 may also be provided at thetail end ofthe fuselage. These surfaces 94 and 95, howr ever, are not controlsurfaces.

The landing gear may be any landing gear of conventional constructionsuch as the pneumatictired wheels 96, at the front of the fuselage,carried by suitable landing gear mechanism such as the lower braces 97and 98 pivotally connected to the fuselage at 99 and the upper brace 100having suitable shock absorbers 101 intermediate the ends thereof andalso suitably pivotally mounted to the fuselage as at 102. A tail skid103 may likewise be provided at the rear end of the fuselage. I

It should be observed that the foregoing, construction may be variedwithout departing from the principles involved in my invention.

Thus, for instance, instead of pivotally supporting the blades 18 uponradial arms 31 and 32, said radial arms may either be replaced oraugmented 41, by a differential setting of the two inner ecby relativelythin tension wires extending taut between the pivotal points of adjacentblades as nected to the rotor controls by trains of gears of relativelylow ratio or mechanical advantage, this construction may be varied tomeet practical requirements, as by greatly increasing the ratio or themechanical advantage between the rotors and the manual controls withinthe fuselage, so as to reduce to a practical degree the amount of manualforce necessary 'to operate the controls and to maintain them in any setposition.

Likewise, if desired, in addition to increasing the ratio of the geartrains, a modified form of mechanism may likewise be employed, by meansof which the controls become non-reversible so' as to relieve theoperator not only of any excessive ,manual force'in operating thecontrols but of all effort in maintaining the controls in any setposition. Thus, by the interposition of suitable worm gears between therotors and the manual controls in the fuselage, a non-reversible effectmay be obtained in the operation of the controls.

In Figures 1 and 2, I have shown the source of power, to wit,the engine90, positioned in the uppermost part of the fuselage with its main shaftaxis intersecting the axis of the driving shaft 19 of the rotors. Inpractice it may be desirable to position the engine 90 in the front orin the lowermost part or near the bottom of the fuselage.

I am aware that the invention may be embodied in other specific formswithout departing from the spirit or essential attributes thereof, and Itherefore desire the present embodiments to be considered in allrespects as illustrative and not restrictive, reference being had to theappended claims rather than to the foregoing description to indicate thescope of the invention.

Having thus described my invention, what I claim as new and desire tosecureby Letters Patent, is: r,

1. In an aircraft, a fuselage, a rotor having its axis of rotationextending'transversely of the direction of travel, said rotor comprisinga pivotally mounted blade, means for oscillating the blade about itspivot between limiting angular positions less than 360 degrees apart,during each revolution of the rotor, means for varying the positions ofthe outer angular limits of blade oscillations in each revolution of therotor, between two predetermined limiting angles-means for positioningthe two limits of angular blade oscillation at any desired part of theorbit of the blades at the will of the operator, and means for varyingthe angle between the two limits of blade oscillation, also at the willof the operator.

3. In an aircraft, a fuselage, a rotor comprising a plurality ofaerofoil blades rotatably mounted about an axis extending transverselyof the direction of travel of the aircraft, a normally stationarycomposite eccentric for automatically deflecting each of the bladesbetween two generally opposed limits of angular deflection, during eachrevolution of the rotor, and means for varying the direction as well asthe magnitude of the effective eccentricity of the composite eccentricat the will of the operator, thereby to vary both the positions of thetwo limits of angular blade deflection in the orbit of the blades andalso to vary the angle between such two opposed limits of bladedeflection. I

4. In an aircraft, a fuselage, a rotor comprising a plurality ofalerofoils arranged to rotateabout an axis extending transversely of thenormal direction of travel of the aircraft, a composite cc"- centriccomprising a plurality of dependent eccentrics, means for independentlyvarying the direction or angle of the eccentricity of each of theplurality of dependent eccentrics at the will of the operator, therebyto vary boththe direction and magnitude of the resultant effective eccentricity of the composite eccentric, and thereby to vary the position ofthe two opposed limits of blade deflection in the orbit of the blades,and also to vary the angle between such limits of blade deflection. i

5. In an aircraft, a fuselage, a rotorfcomprising a plurality ofaerofoil blades adapted to revolve generally about an axis extendingtransversely of the normal direction of travel of the aircraft, a.normally stationary, composite eccentric, means intermediate saidcomposite eccentric and each" of said blades for angularly oscillatingeach of said blades between two opposed and predetermined limitingpositions during each revolution of the rotor, said composite eccentriccomprising a plurality of normally stationary and adjustable, dependenteccentrics, and means for independently varying the direction of theeccentricity of each of said dependent eccentrics "at the will of theoperator, thereby to vary both the direction as well as the magnitude ofthe resultant effective eccentricity of the composite eccentric, at thewill of the operator.

6. In an aircraft, a fuselage, a pair of opposed rotors, each includinga plurality of aerofoil blades adapted to revolve generally about anaxis extending transversely of the-direction of travel of the aircraft,automatic means for angularly oscillating each of the blades of therotors between two opposed limits during each revolution of said rotors,means for varying the positions of the two opposed limits of angularblade deflection in the orbit of the blades at the will of the operator,and means for varying the angle between the two limits of angular bladedeflection at the will of the operator 7. In an aircraft, a fuselage, apair of opposed rotors, each including a plurality of aerofoil bladesadapted to revolve generally about an axis extending transversely of thenormal direction of travel of the aircraft, automatic means forangularly oscillating each of the blades of the rotors between twoopposed limits during each revolution of said rotors, means for varyingthe positions of the two opposed limits of angular blade I deflection inthe orbit of the blades at the will of intermediate the compositeeccentric and each of the blades of the rotors for automaticallyoscillating each of the blades between two opposed limits of angularblade deflection during each revolution of the rotors, each of saidcomposite eccentrics comprising a plurality of dependent eccentrics, andmeans for varying the direction of eccentricity of each of saiddependent eccentrics comprising each of the composite eccentrics. eitherjointly or independently of each other at the will of the operator,thereby to vary either the direction or the magnitude of the resultanteffective eccentricity of the composite eccentrics or to vary both thedirection and the magnitude of the effective eccentricity of saidcomposite eccentrics.

9. In an aircraft, a fuselage, a pair of opposed rotors, each comprisinga plurality of aerofoil blades adapted to revolve generally about anaxis extending transversely of the normal direction of the travel of theaircraft, a normally stationary, composite eccentric in each of saidrotors, means intermediate the composite eccentric and each of theblades of the rotors for automatically oscillating each of the bladesbetween two opposed limits of angular blade deflection during eachrevolution of the rotors, each of said composite eccentrics comprising aplurality of dependent eccentrics, means for varying the direction ofeccentricity of each of said dependent I composite eccentrics, and meansfor effecting such joint or independent variations of directions ofeccentricities in each of the pair of opposed rotors, either uniformlyor differentially, at the will of the operator.

10. In an aircraft, a fuselage, a rotor comprising a plurality ofpivotally mounted aerofoil blades adapted to travel generally about acommon axis and extending transversely of the normal direction of travelof the aircraft, a source of power for revolving said rotor, means foroscillating each of said blades in succession during porting frame, aneccentric carried by said stationary frame member, means intermediatesaid eccentric and each of said blades, having pivotal connection withthe latter, manual controls within the fuselage, and means intermediatesaid manual controls and said eccentric for varying the direction andthe magnitude of the eccentricity thereof, at the will of the operator.

12. In an aircraft, a fuselage, a driving shaft, a rotor carried by saiddriving shaft comprising a plurality of blades, means intermediate saiddriving shaft and each of said blades, rigidly secured to the former andhaving pivotal supporting connection with the latter at substantiallythe center of gravity of the cross section thereof,

' a stationary supporting frame, an eccentric carried by said stationaryframe member, means intermediate said eccentric andeach of said blades,having pivotal connection with the latter, manual controls within thefuselage and means intermediate said manual controls and said eccentricfor varying the direction and the magnitude of the eccentricity thereof,at the will bf the operator.

13, In an aircraft, afuselage, a driving shaft, a rotor carried by saiddriving shaft comprising a plurality of blades, means intermediate saiddriving shaft and each of said blades, rigidly secured to the former andhaving pivotal supporting connection with the latter, a stationarysupporting frame, an eccentric carried by said sta tionary frame member,means intermediate said eccentric and each of said blades, havingpivotal connection with the latter, in advance of the pivotal supportingconn'ection thereof, manual controls within the fuselage, and meansintermediate said manual controls and said eccentric for varying thedirection and'the magnitude of the eccentricity thereof, at the will ofthe operator.

14. In an aircraft, a fuselage, a driving shaft, a rotor carried by saiddriving shaft comprising a plurality of blades, means intermediate saiddriving shaft and each of said blades, rigidly secured to the former andhaving pivotal supporting connection with the latter, at substantiallythe center of gravity of the cross section thereof,

.a stationary supporting frame, an eccentric carried by said stationaryframe member, means intermediate said eccentric and each of said blades,having pivotal connection with the latter, in advance of the pivotalsupporting con-, nection thereof, manual controls within the fuselageand means intermediate said manual controls and said eccentric forvarying the direction and the magnitude of the eccentricity thereof, atthe will of the operator.

15. In an aircraft, a fuselage, a driving shaft, a rotor carried by saiddriving shaft comprising a plurality of blades, means intermediate saiddriving shaft and each of said blades, rigidly secured to the former andhaving pivotal supporting connection with the latter, a stationarysupporting frame, a normally stationary eccentric carried by saidstationaryframe member, a second normally stationary eccentric carriedby said first eccentric, means intermediate said second eccentric andeach of said blades, having pivotal connection with the latter, andmeans for independently rotating each of said eccentrics.

16. In an aircraft, a fuselage, a driving shaft, a driving spiderrigidly attached to said driving shaft, a plurality of blades pivotallymounted on said driving spider, a stationary supporting frame, anormally stationary eccentric carried by said supporting frame, a secondnormally stationary eccentric carried by said first eccentric, aconnecting rod intermediate said second eccentric and each of saidblades, and means for independently rotating each of said eccentrics.

1'7. An aircraft, comprising a fuselage, a rotor, comprising a pluralityof pivotally mounted aerofoil blades, extending generally transverselyto the direction of travel of the aircraft and adapted to revolve in anorbit about the axis of the rotor, means for automatically oscillatingeach of said aerofoil blades in succession about its respective pivotduring each revolution of the rotor, and means for varying the magnitudeof the oscillation and the event of the oscillation in the travel of theaerofoil blades about the axis of the rotor, at the will of theoperator.

18. An aircraft comprising a-fuselage, a rotor comprising a plurality ofpivotally mounted aerofoil blades extending generally transversely tothe direction of travel of the aircraft and adapted 19. An aircraftcomprising a fuselage, a rotor comprising a shaft and a plurality ofpivotally mounted aerofoil blades carried by said shaft and being spacedtherefrom generally equidistantly, and having cambered transversecroa-sections, said shaft and blades extending generally transversely ofthe normal direction of travel of the aircraft and said shaft beingdisposed generally above the center of gravity of the fuselage; meansfor automatically oscillating, between two predetermined limitingangles, each of said blades, in succession, about its respective pivots,during each traverse through the orbit of said' blades, means forvarying the limits of oscillation of said blades with relation to eachother and with relation to the orbit of said blades, and a prime moveroperatively connected to said rotor, rotatably to propel the same in adirection such that when the rotor is power-driven, the torque reactionof the rotor upon the fuselage will tend to raise the rear end of thefuselage,

20. In an aircraft, a fuselage, a rotor having its axis of rotationextending transversely of the direction of travel, said rotor comprisinga piv tally mounted aerofoil blade having a cambered transversecross-section, means for oscillating the blade about its pivot betweenlimiting angular ra se positions less than :00 degrees apart, duringeachrevolution of the rotor, means forvaryins the positions of the outerangular limits of blade oscillations in the orbit oi the blade, andmeans for varying the maximum angle of deflection of the blade in itsorbit about the axis of the rotor.

21. In an aircraft, a fuselage. a rotor comprising a plurality ofpivotally mounted aerofoil blades having 'cambered transversecross-sections, adapted to travel in an orbit about an axis extendingtransversely of the direction of travel of,

the aircraft, means for automatically oscillating about its pivot eachof the blades in succession during each revolution of the rotor, betweentwo.

predetermined limiting angles, means for positioning the two limits ofangular blade oscillations at anydesired part of the orbit of the bladesat the will of the operator, and meansforvarying the angle between thetwo limits of blade oscillatlon, also at the will of the operato 22. Inan aircraft, a fuselage. a rotor compr'is- I ing a plurality of aerofoilblades having cambe'red fective eccentricity of the composite eccentricatthe willof the operator, thereby to vary both the positions of the twolimits of angular blade deflection in the orbit of the blades and alsoto vary the angle between such two opposed limits of blade deflection.

navrmnn n. PLA'I'I.

transverse cross-sections, rotatablymountedabout

